Fan blade anti-icing concept

ABSTRACT

A fan blade anti-icing system comprises a fan hub and a fan blade extending radially outwardly from the fan hub. The fan blade has a base and an airfoil extending radially outwardly from the fan base. The airfoil having a leading edge, a trailing edge, a convex side surface between the leading and trailing edge and a concave side surface between the leading and trailing edge. The fan blade further has a radial passage extending from a blade air inlet in the blade base in communication with a source of heated air, and a rearwardly directed passage in communication with the radial passage and having a blade air outlet forward of the trailing edge and oriented tangentially to the convex side surface or concave side surface of the airfoil.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a divisional of U.S. application Ser. No. 16/683,552filed Nov. 14, 2019, which claims priority to U.S. provisional patentapplication No. 62/925,848 filed Oct. 25, 2019, the entire contents ofwhich are incorporated by reference herein.

TECHNICAL FIELD

The disclosure relates generally to anti-icing of fan blades in a gasturbine engine.

BACKGROUND

Ice can form and adhere to the fan blades of a gas turbine engine undercertain conditions during flight. The weight of the ice buildup canresult in imbalance of the fan and can be detrimental to efficient airflow. When ice breaks away and is released, the fan can be imbalanced,orbiting increases, vibration occurs and impact from ice particles cancause foreign object damage.

SUMMARY

The disclosure describes a fan blade anti-icing system for a gas turbineengine comprising: a fan hub mounted for rotation about an axis; and afan blade extending radially outwardly from the fan hub, the fan bladehaving a base and an airfoil extending radially outwardly from the fanbase, the airfoil having a leading edge, a trailing edge, a convex sidesurface between the leading and trailing edge and a concave side surfacebetween the leading and trailing edge, the fan blade further having aradial passage extending from a blade air inlet in the blade base incommunication with a source of heated air, and a rearwardly directedpassage in communication with the radial passage and having a blade airoutlet upstream of the trailing edge and oriented tangentially to theconvex side surface or concave side surface of the airfoil.

In accordance with another aspect, there is provided a fan blade for agas turbine engine comprising: a fan blade having a blade base and anairfoil with a radially outward axis, the airfoil having a leading edge,a trailing edge, a convex side surface between the leading and trailingedge and a concave side surface between the leading and trailing edge,the fan blade further having a radial passage extending from a blade airinlet in the blade base for communication with a source of heated air, arearwardly directed passage in communication with the radial passage andhaving a blade air outlet upstream of the trailing edge and orientedtangentially to the convex side surface or the concave side surface ofthe airfoil.

In accordance with a still further general aspect, there is provided amethod of impeding icing on an airfoil surface of a fan blade of anaircraft engine, the method comprising: receiving heated pressurized airinside the fan blade; and directing the heated pressurized air exitingthe fan blade to flow in a downstream direction over the airfoil surfaceof the fan blade.

Further details of these and other aspects of the subject matter of thisapplication will be apparent from the detailed description includedbelow and the drawings.

DESCRIPTION OF THE DRAWINGS

FIG. 1 shows an axial cross-section view of a turbo-fan gas turbineengine.

FIG. 2 is an isometric transparent view of a section through anintegrally bladed fan rotor in accordance with the present description.

FIG. 3 is an enlarged view of FIG. 2 showing an internal radial passageand multiple rearwardly directed passages with heated air outlets tocreate a Coanda effect heated air film over the airfoil surface of a fanblade.

FIG. 4 is a further isometric transparent view of the fan blade as shownin FIG. 3 with a radial passage and having multiple rearwardly directedpassages.

FIG. 5 is an isometric solid view of an alternative fan blade having arecessed pocket and cover defining an air plenum, the cover includingmultiple outlets to provide for the formation of a film of anti-icingair over a surface of the airfoil.

FIG. 6 is an isometric solid view along line 6-6 of FIG. 5 showing thedrilled radial passage.

FIG. 7 is a view of the recessed pocket of FIG. 5 with the coverremoved.

FIG. 8 is a radial section along line 8-8 of FIG. 7 showing theintersection of the radial passage with the recessed pocket.

FIG. 9 is a further alternative to FIG. 5 with the multiple outlets in aradial alignment.

DETAILED DESCRIPTION

FIG. 1 shows an axial cross-section through an aircraft engine.According to the illustrated embodiment, the aircraft engine is aturbo-fan gas turbine engine. It is understood that the aircraft enginecould adopt various forms others than the illustrated example. Airintake into the engine passes over fan blades 1 in a fan case 2 and isthen split into an outer annular flow through the bypass duct 3 and aninner flow through the low-pressure axial compressor 4 and high-pressurecentrifugal compressor 5. Compressed air exits the compressor through adiffuser 6 and is contained within a plenum 7 that surrounds thecombustor 8. Fuel is supplied to the combustor 8 through fuel tubes 9and fuel is mixed with air from the plenum 7 when sprayed throughnozzles into the combustor 8 as a fuel air mixture that is ignited. Aportion of the compressed air within the plenum 7 is admitted into thecombustor 8 through orifices in the side walls to create a cooling aircurtain along the combustor walls or is used for cooling to eventuallymix with the hot gases from the combustor and pass over the nozzle guidevane 10 and turbines 11 before exiting the tail of the engine asexhaust.

The present description and drawings relate to anti-icing features ofthe fan blades 1. The compressors 4, 5 and combustor 8 createpressurized air having a temperature greater than ambient and at leastabove the freezing temperature of water at flight altitude. Heatedpressurized air can be bled from the compressors 4, 5 and combustor 8and directed through the engine to the fan hub 12 via ducts within thehollow central engine shafts for example.

With reference to FIGS. 2 and 3, there is shown an integrally bladed fanrotor having a fan hub 12 with a hub air inlet in communication with asource of heated air, such as the compressors 4, 5 and combustor 8. Theexemplified fan hub 12 is generally cylindrical or conical with aradially outer surface with multiple blades spaced about thecircumference. Each blade has a blade base 14. The fan hub 1 has aplurality of hub air outlets in communication with the hub air inlet todistribute heated pressurized air to each blade.

The fan blade has an airfoil 15 with a radially outward axis 16generally normal to the direction of air flow into the engine (seearrow). The airfoil 15 has a leading edge 17, a trailing edge 18, aconvex side surface 19 between the leading and trailing edge 17, 18 anda concave side surface 20 between the leading and trailing edge 17, 18.

As best seen in FIGS. 3 and 4, the blade has a radial passage 21 forreceiving and distributing heated pressurized air. The radial passage 21extends from a blade air inlet 22 in the blade base 14 that is incommunication with the air outlet in the fan hub 12. The radial passage21 is provided adjacent to the leading edge 17. In the exampleillustrated, six rearwardly directed passages 23 branch off in fluidcommunication with the radial passage 21. Each passage 23 extendstowards the trailing edge 18 and terminates in a blade air outlet 24 inthe concave side surface 20, the convex side surface 20 or both.

The blade air outlets 24 are disposed upstream of the trailing edge 18typically within the upstream half of the airfoil and orientedpredominantly tangential to the airfoil surface to emit heatedpressurized air substantially parallel to the incoming air that passesover the airfoil 15. The rearward passages 23 and blade air outlets 24are oriented in a rearward or downstream direction substantiallyparallel to the incoming air direction to provide for the formation of aheated air film over the airfoil surface. In the radially inward area ofthe airfoil 15 adjacent to the blade platform 25 and fillet 26, theincoming air direction is directed to be parallel to the blade platform25 (see arrow in FIG. 2).

The heated pressurized air from the blade air outlets 24 is ejected instreams or jets that merge smoothly with the incoming air. As a result,the streams of heated air are subjected to the Coanda effect and flowdownstream attached to the concave side surface 20. The Coanda effect isthe tendency of a fluid jet to stay attached to an adjacent surface,named after Romanian inventor Henri Coanda. The effect is the tendencyof a jet of fluid emerging from an orifice to follow an adjacent flat orcurved surface and to draw in or entrain fluid from the surroundings sothat a region of lower pressure develops. The lower pressure regionbetween the jet and adjacent surface draws the jets towards the adjacentsurface to flow parallel to or “attach” to the surface. Eventually thejet and ambient air flow tend to mix downstream due to turbulence andthe Coanda effect dissipates.

Therefore, the jet of heated pressurized air from the blade air outlets24 creates a film of heated air flowing parallel to and closely attachedto the airfoil surface. The heated air flowing through the radialpassage 21 and the multiple rearward passages 23 will heat the metal ofthe airfoil 15 through convection. In addition, the rearward orientationof the rearward passages 23 and blade air outlets 24 will emit jets ofheated air that will flow close to and parallel to the concave sidesurface 20, the convex side surface 19 or both resulting from the Coandaeffect. The areas of the airfoil surface over which the heated jets ofair flow will be locally heated to impede formation of ice and melt icethat has been deposited.

In the illustrated example, the multiple rearwardly directed passages 23and blade air outlets 24 are radially spaced apart and are radiallyaligned on an imaginary line that is transverse to the blade platform25. Various alternative patterns of location the rearwardly directedpassages 23 and blade air outlets 24 can be adopted depending on thenature of blade anti-icing required. In the examples illustrated, asseen in FIG. 4 the blade air outlets 24 are disposed in an area of theconcave side surface 20 that is an offset by a distance ‘x’ from theblade platform 25.

FIGS. 5 to 9 show an alternative fan blade anti-icing arrangement with asingle rearwardly directed passage formed as a recessed pocket 27 (seeFIG. 7) in the concave side surface 28. As best seen in FIG. 7, therecessed pocket 27 intersects with and is in fluid communication withthe radial passage 29. FIG. 6 shows the blade air inlet 30 and radialpassage 29 that have been drilled to connect with the recessed pocket27.

As seen in FIG. 5, the recessed pocket 27 is covered with a cover 31 todefine an internal plenum within the airfoil 32. The plenum between thecover 31 and the recessed pocket 27 receives heated pressurized air fromthe radial passage 29. The cover 31 and adjacent areas of the airfoil 32are heated by convection to locally impede icing. Heated air is emittedthrough the blade air outlets 33 that fluidly communicate with theinternal plenum beneath the cover 31. Heated air from the blade airoutlets 33 is emitted rearwardly in streams or jets that aresubstantially parallel to the blade platform 34 and substantiallyparallel to the rearward incoming air flow direction (see arrow).

FIGS. 5 and 9 show blade air outlets 33 and blade air outlets 35 drilledinto the cover 31 and cover 36. The blade air outlets 33, 35 can bedrilled through the cover 31 and 36, or may be formed as a slot in anedge of the cover 31, 36 or may be formed as a slot in an edge of therecessed pocket 27. The cover 31, 36 may be sealed to the edges of therecessed pocket 27 by adhesive bonding, welding, or diffusion bondingfor example. The corners of the recessed pocket 27 and cover 31 arerounded to reduce stress concentration issues. The cover 31, 36 could beformed as a round disk or other shapes if desired. The shape andlocation of the recessed pocket 27, and the arrangement of blade airoutlets 33, 35 can be selected to target the specific icing experiencedby any blade configuration. The above described system can be applied tobladed and integrally bladed fans, both solid and hollow blades, andboth metal and composite blades.

The above description is meant to be exemplary only, and one skilled inthe relevant arts will recognize that changes may be made to theembodiments described without departing from the scope of the inventiondisclosed. The present disclosure may be embodied in other specificforms without departing from the subject matter of the claims. Thepresent disclosure is intended to cover and embrace all suitable changesin technology. Modifications which fall within the scope of the presentinvention will be apparent to those skilled in the art, in light of areview of this disclosure, and such modifications are intended to fallwithin the appended claims. Also, the scope of the claims should not belimited by the preferred embodiments set forth in the examples, butshould be given the broadest interpretation consistent with thedescription as a whole.

What is claimed is:
 1. A method of impeding icing on an airfoil surfaceof a fan blade of an aircraft engine, the airfoil surface having aconcave side surface and a convex side surface opposite to the concaveside surface, the method comprising: receiving heated pressurized airinside the fan blade; ejecting the heated pressurized air adjacent to aleading edge thereof; and using a Coanda effect to flow the heatedpressurized air over the airfoil surface towards a trailing edge of thefan blade, wherein the heated pressurized air is ejected on both theconvex and concave side surfaces.
 2. The method of claim 1, wherein theheated pressurized air exiting the fan blade is ejected tangentially tothe airfoil surface.
 3. The method of claim 1, wherein the heatedpressurized air is ejected in a radially inward area of the airfoilsurface adjacent to a blade platform of the fan blade.
 4. The method ofclaim 1, wherein the heated pressurized air received inside the fanblade includes air bled from a compressor the aircraft engine.
 5. Themethod of claim 1, wherein the heated pressurized air is ejected in jetstangentially to incoming air flowing over the fan blade, and wherein thejets are spaced-apart along a spanwise direction of the fan blade.
 6. Amethod of mitigating ice accretion on an airfoil surface of a fan bladeof an aircraft engine, the airfoil surface having a concave side surfaceand a convex side surface opposite to the concave side surface, themethod comprising: receiving heated pressurized air inside the fanblade; ejecting the heated pressurized air within an upstream half ofthe airfoil and in a downstream direction predominantly tangential tothe airfoil surface and parallel to incoming air that passes over theairfoil surface of the fan blade during engine operation; and using theheated pressurized air ejected from the fan blade to create a film ofheated air flowing in a downstream direction over the airfoil surfacetowards a trailing edge of the fan blade, wherein the heated pressurizedair is ejected on both the convex and concave side surfaces.
 7. Themethod of claim 6, wherein the heated pressurized air is ejectedadjacent to a leading edge of the fan blade.
 8. The method of claim 7,wherein the heated pressurized air is ejected in a radially inward areaof the airfoil surface adjacent to a blade platform of the fan blade. 9.The method of claim 8, comprising using a Coanda effect to create thefilm of heated air over the airfoil surface.
 10. The method of claim 9,wherein the heated pressurized air is ejected in jets tangentially tothe incoming air flowing over the fan blade, and wherein the jets arespaced-apart along a spanwise direction of the fan blade.